The present invention relates generally to rotor blades (such as those used in turbines, compressors, fans, and the like in a gas turbine engine), and more particularly to such a rotor blade having improved internal cooling.
Gas turbine engines, such as aircraft jet engines, include turbines having rotor blades. A turbine rotor blade has a shank which is attached to a rotating turbine rotor disk and an airfoil blade which is employed to extract useful work from the hot gasses exiting the engine's combustor. The airfoil blade includes a blade root which is attached to the shank and a blade tip which is the free end of the airfoil blade. Modern aircraft jet engines have employed internal cooling of turbine rotor blades to keep the airfoil blade temperatures within design limits. Typically, the airfoil blade portion of the turbine rotor blade is cooled by air (typically bled from the engine's compressor) passing through a longitudinally extending cylindrical internal passage, with the air entering near the airfoil blade root and exiting near the airfoil blade tip. Known turbine blade cooling passages include a cooling circuit comprising a plurality of unconnected longitudinally-oriented passages each receiving cooling air from near the airfoil blade root and channeling the air longitudinally toward the airfoil blade tip. Other known cooling circuits include a serpentine cooling circuit comprising a plurality of longitudinally-oriented passages which are series-connected to produce serpentine flow. For either cooling circuit, some air also exits the airfoil blade through film cooling holes near the airfoil blade's leading edge, and some air exits the airfoil blade through trailing edge cooling holes.
Cooling passages typically have circular, rectangular, square or oblong transverse cross-sectional shapes. It is known that for a rotating airfoil blade having a serpentine cooling circuit including longitudinally-oriented cooling passages of square cross-sectional shape, Coriolis (rotation) forces will increase the heat transfer coefficient (by a factor of more than two in one reported experiment) along certain walls of the passage and decrease the heat transfer coefficient (by a factor of more than two in the reported experiment) along other walls of the passage as compared with a non-rotating airfoil. Basically, the Coriolis force is proportional to the vector cross product of the velocity vector of the coolant flowing through the passage and the angular velocity vector of the rotating airfoil blade. The Coriolis force compresses the coolant against one side of the square passage increasing the heat transfer at that side while decreasing the heat transfer at the opposite side. This creates an uneven transverse cross section blade temperature profile which creates hot areas that must be compensated for by, for example, increasing the cooling flow. Increasing the cooling flow could be accomplished by bleeding off more engine compressor air, but this would reduce the engine's efficiency by reducing the number of miles flown for each gallon of fuel consumed. What is needed is a cooling passage configuration which takes advantage of, rather than suffers from, the effects of Coriolis forces on the effectiveness of transferring heat from the airfoil blade to the coolant.